This invention relates generally to cooling of airfoils and, more particularly, to a method and apparatus for cooling the trailing edges of gas turbine airfoils.
A well developed field exists regarding the investment casting of internally-cooled turbine engine parts such as blades and vanes. In an exemplary process, a mold is prepared having one or more mold cavities, each having a shape generally corresponding to the part to be cast. An exemplary process for preparing the mold involves the use of one or more wax patterns of the part. The patterns are formed by molding wax over ceramic cores generally corresponding to positives of the cooling passages within the parts. In a shelling process, a ceramic shell is formed around one or more such patterns in well known fashion. The wax may be removed such as by melting in an autoclave. This leaves the mold comprising the shell having one or more part-defining compartments which, in turn, contain the ceramic core(s) defining the cooling passages. Molten alloy may then be introduced to the mold to cast the part(s). Upon cooling and solidifying of the alloy, the shell and core may be mechanically and/or chemically removed from the molded part(s). The part(s) can then be machined and treated in one or more stages.
The ceramic cores themselves may be formed by molding a mixture of ceramic powder and binder material by injecting the mixture into hardened steel dies. After removal from the dies, the green cores are thermally post-processed to remove the binder and fired to sinter the ceramic powder together. The trend toward finer cooling features has taxed core manufacturing techniques. The fine features may be difficult to manufacture and/or, once manufactured, may prove fragile. Commonly-assigned co-pending U.S. Pat. No. 6,637,500 of Shah et al. discloses general use of a ceramic and refractory metal core combination. There remains room for further improvement in such cores and their manufacturing techniques.
The currently used ceramic cores limit casting designs because of their fragility and because cores with thickness dimensions of less than about 0.012-0.015 inches cannot currently be produced with acceptable casting yields.
The trailing edge cut-back geometry is one of the most utilized cooling configurations in airfoil design. This preferred application stems from two practical standpoints. First, the aerodynamic losses associated with such a blade attain the lowest values due to a thinner trailing edge. Second, airfoil high pressure side heat load to the part is reduced by using film cooling at the trailing edge pressure side.
Smaller trailing edge thickness leads to a lower pressure difference between the pressure and the suction sides of the airfoil. Trailing edge configurations without cut-back, known as centerline cooling tailing edges, with a pressure-to-suction side pressure ratio of about 1.35, results in trailing edge thickness in the order of 0.050 in. For these centerline discharge designs, the total pressure loss at 50 percent radial span could be as high as 3.75 percent. This relatively high pressure loss leads to undesirable high aerodynamic losses. A practical way to reduce these losses, is to use a pressure side ejection trialing edge configuration with a cut-back length. In such a configuration, the trailing edge can attain a thickness as low as 0.030 in. to reduce the aerodynamic losses. Typical of such a cut-back design is that shown in U.S. Pat. No. 4,601,638, assigned to the assignee of the present invention and incorporated herein by reference.
In this context, there are several internal cooling design features that control the heat transfer at the trailing edge. These can be summarized as follows: (1) size of the cooling passage; (2) internal cooling features inside the cooling passage; (3) trailing edge thickness distributions; (4) pressure side trailing edge lip thickness; (5) pressure side land roughness, and (6) slot film cooling coverage. It should be noted that only elements (1) and (2) can be used effectively for centerline discharge tailing edge designs; whereas all elements (1) through (6) can be used for the pressure side ejection design with a cut-back trailing edge. In the pressure side ejection designs, the thermal-mechanical fatigue and creep life will also improve with improved metal temperature distributions for the entire trailing edge region.
In general, the external thermal load on the airfoil pressure side is about two times that of the suction side, and therefore, there is a greater potential for pressure side fatigue to occur on the airfoil pressure side. Under cyclic conditions, crack nucleation may also occur sooner on the pressure side.
Since the airfoil trailing edge responds faster than the rest of the airfoil due to its lower thermal mass; these areas are particularly prone to fatigue failure. Crack nucleation leads to linkage with thermal-mechanical fatigue cracking, originating and propagating from the trailing edge. As cracks propagate, load shakedown will occur throughout the blade as the load is redistributed to other portions of the trailing edge. This is particularly true for rotating blades as the centrifugal load remains constant. Load shakedown leads to overload conditions, or conditions where the stresses in the blade may be above yield stress of the material as the load bearing blade area has decreased due to cracking. The material will start deforming plastically even at colder parts of the airfoil. This is an irreversible effect leading in all likelihood to blade liberation and failure. Thus, selection of the trailing edge pressure side ejection design for cooling a blade trailing edge region becomes crucial.
At the trailing edge regions, internal impingement configurations have been used in the gas turbine airfoil design. In general, cooling air is allowed to pass through rib cross-over openings leading to jet impingement onto subsequent ribs and surrounding walls. The flow acceleration is high through these cross-over impingement openings. The coolant flow Mach number profile follows that of the coolant static pressure profile in that it assumes an almost step-wise profile at these openings. The step-wise profiles are undesirable as they lead to relatively high peaks in internal heat transfer coefficients at the walls of the blade. In other words, there are regions in the airfoil trailing edge wall, which attain areas of relatively lower metal temperatures with high internal heat transfer coefficients. Meanwhile, other areas with lower internal convective heat transfer coefficients lead to relatively higher metal temperatures. These metal temperature differences lead to high thermal strains, which in conjunction with transient thermal stresses in the airfoil during take-off, in turn, lead to undesirable thermal-mechanical fatigue problems in the airfoil trailing edge.